The present invention relates generally to gas turbine engines, and, more specifically, to turbine cooling therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Multiple turbine stages follow the combustor for extracting energy from the combustion gases to power the compressor and produce useful work.
In a typical turbofan gas turbine engine configuration, a high pressure turbine (HPT) immediately follows the combustor for receiving the hottest combustion gases therefrom from which energy is extracted for powering the compressor. A low pressure turbine (LPT) follows the HPT and extracts additional energy from the combustion gases for powering a fan disposed upstream from the compressor for producing propulsion thrust for powering an aircraft in flight.
The HPT includes a turbine nozzle at the discharge end of the combustor which directs the combustion gases between first stage turbine rotor blades arranged in a row around the perimeter of a supporting rotor disk. The disk in turn is joined by a corresponding shaft to the rotor of the compressor for rotating the corresponding compressor blades therein.
The nozzle vanes and rotor blades have corresponding airfoil configurations specifically tailored for maximizing energy extraction from the hot combustion gases. The vanes and blades are hollow and include internal cooling circuits which typically use a portion of the compressor discharge pressure (CDP) air for cooling thereof. Since the nozzle vanes are stationary and the rotor blades rotate during operation, they typically have different internal cooling configurations, while similarly sharing various rows of film cooling holes through the pressure and suction sides thereof for providing external film cooling of the vanes and blades.
Any CDP air diverted from the combustion process decreases efficiency of the engine and should be minimized. However, sufficient cooling air must be used to limit the operating temperature of the vanes and blades for ensuring a suitable useful life thereof.
The turbine vanes and blades are typically manufactured from state-of-the-art superalloy materials, typically nickel or cobalt based, which have high strength at the elevated temperatures experienced in a modern gas turbine engine. The use of superalloy material and intricate cooling circuits in turbine vanes and blades helps minimize the requirement for diverting discharge air from the compressor for cooling thereof.
Furthermore, typical commercial aircraft have well defined operating cycles including takeoff, cruise, descent, and landing, with the engine being operated with a correspondingly short duration at maximum power or high turbine rotor inlet temperature.
In the continuing development of advanced gas turbine engines, it is desirable to operate the engine almost continuously at very high compressor discharge temperature and at correspondingly high turbine rotor inlet temperatures for extended periods of time for maximizing efficiency or performance. This type of engine may be used to advantage in small business jets or advanced military applications.
However, this long and hot operating condition presents extreme challenges in cooling the high pressure turbine rotor using the currently available superalloy disk materials. By operating the compressor for achieving high discharge pressure of the air used in the combustion process, the temperature of that high pressure air is correspondingly increased which decreases the ability of that CDP air to cool the high pressure turbine. Adequate cooling of the turbine is required for ensuring a long useful life thereof and reduce the need for periodic maintenance.
Accordingly, it is desired to provide a gas turbine engine having an improved cooling configuration for the high pressure turbine thereof.